EP4524384A1 - Ensemble moteur pour système de propulsion d'aéronef - Google Patents

Ensemble moteur pour système de propulsion d'aéronef Download PDF

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Publication number
EP4524384A1
EP4524384A1 EP24200082.6A EP24200082A EP4524384A1 EP 4524384 A1 EP4524384 A1 EP 4524384A1 EP 24200082 A EP24200082 A EP 24200082A EP 4524384 A1 EP4524384 A1 EP 4524384A1
Authority
EP
European Patent Office
Prior art keywords
assembly
engine
electric motor
rotational
bladed
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
EP24200082.6A
Other languages
German (de)
English (en)
Inventor
Xi Wang
Jean Thomassin
David Gagnon-Martin
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Pratt and Whitney Canada Corp
Original Assignee
Pratt and Whitney Canada Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Pratt and Whitney Canada Corp filed Critical Pratt and Whitney Canada Corp
Publication of EP4524384A1 publication Critical patent/EP4524384A1/fr
Pending legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/30—Aircraft characterised by electric power plants
    • B64D27/33—Hybrid electric aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/04—Aircraft characterised by the type or position of power plants of piston type
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/10—Aircraft characterised by the type or position of power plants of gas-turbine type 
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D31/00—Power plant control systems; Arrangement of power plant control systems in aircraft
    • B64D31/16—Power plant control systems; Arrangement of power plant control systems in aircraft for electric power plants
    • B64D31/18—Power plant control systems; Arrangement of power plant control systems in aircraft for electric power plants for hybrid-electric power plants
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D35/00—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions
    • B64D35/02—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions specially adapted for specific power plants
    • B64D35/021—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions specially adapted for specific power plants for electric power plants
    • B64D35/022—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions specially adapted for specific power plants for electric power plants of hybrid-electric type
    • B64D35/023—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions specially adapted for specific power plants for electric power plants of hybrid-electric type of series-parallel type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
    • F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/20—Adaptations of gas-turbine plants for driving vehicles
    • F02C6/206—Adaptations of gas-turbine plants for driving vehicles the vehicles being airscrew driven
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K5/00—Plants including an engine, other than a gas turbine, driving a compressor or a ducted fan
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/026—Aircraft characterised by the type or position of power plants comprising different types of power plants, e.g. combination of a piston engine and a gas-turbine
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/70—Application in combination with
    • F05D2220/76—Application in combination with an electrical generator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/40—Transmission of power
    • F05D2260/403—Transmission of power through the shape of the drive components
    • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
    • F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type

Definitions

  • This invention relates generally to an engine assembly for an aircraft propulsion system and, more particularly, to an engine assembly including an intermittent internal combustion engine and a gas turbine engine configured to cooperatively drive a propulsor of the aircraft propulsion system.
  • a propulsion system for an aircraft may be configured, for example, with an engine assembly including an intermittent internal combustion engine such as a rotary engine.
  • an engine assembly including an intermittent internal combustion engine such as a rotary engine.
  • Various configurations of such engine assemblies are known in the art. While these known engine assemblies have various advantages, there is still room in the art for improvement.
  • an engine assembly for an aircraft propulsion system includes a propulsor, a gearbox, an engine, a gas turbine engine, and an electric control assembly.
  • the gearbox includes a gear assembly coupled with the propulsor.
  • the engine includes an air inlet, an exhaust outlet, and an engine output shaft.
  • the engine output shaft is coupled with the gear assembly and configured to drive rotation of the propulsor through the gear assembly.
  • the gas turbine engine includes a first rotational assembly, a compressor section, a turbine section, and a combustor.
  • the first rotational assembly is configured for rotation about a rotational axis.
  • the first rotational assembly includes a bladed compressor rotor for the compressor section, a bladed first turbine rotor for the turbine section, and a first shaft interconnecting the bladed compressor rotor and the bladed first turbine rotor.
  • the compressor section is connected to the air inlet and configured to direct a compressed air to the air inlet.
  • the combustor is connected to the exhaust outlet and configured to receive an exhaust gas from the exhaust outlet.
  • the combustor is configured to direct a combustion gas through the turbine section to drive rotation of the bladed first turbine rotor.
  • the electric control assembly includes a first electric motor and an energy storage device.
  • the first electric motor is coupled with the gear assembly.
  • the first electric motor is electrically connected to the energy storage device.
  • the first electric motor is configured to selectively apply a rotational force to the gear assembly to further drive rotation of the propulsor through the gear assembly.
  • the gas turbine engine may further include a second rotational assembly.
  • the second rotational assembly may include a bladed second turbine rotor for the turbine section and a second shaft.
  • the second shaft may interconnect the bladed second turbine rotor and the gear assembly to further drive rotation of the propulsor through the gear assembly.
  • the combustor may be configured to direct the combustion gas through the turbine section to drive rotation of the bladed first turbine rotor and the bladed second turbine rotor.
  • the second rotational assembly may be axially separated from the first rotational assembly.
  • the bladed second turbine rotor may be downstream of the first bladed turbine rotor with respect to the combustion gas.
  • the electric control assembly may further include a second electric motor.
  • the second electric motor may be coupled with the first rotational assembly.
  • the second electric motor may be electrically connected to the energy storage device.
  • the engine assembly may further include a controller including a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, may cause the processor to control the second electric motor to generate a first electrical power during a first operating condition of the engine assembly.
  • the instructions when executed by the processor, may further cause the processor to control the first electric motor to apply the rotational force to the gear assembly during the first operating condition of the engine assembly while the second electric motor generates the first electrical power.
  • the instructions when executed by the processor, may further cause the processor to control the second electric motor to generate the first electric power during the first operating condition of the engine assembly to charge the energy storage device.
  • the instructions when executed by the processor, may further cause the processor to control the first electric motor to generate a second electrical power during a second operating condition of the engine assembly.
  • the second operating condition may be different than the first operating condition.
  • the engine may be an intermittent internal combustion engine.
  • an engine assembly for an aircraft propulsion system includes a propulsor, a gearbox, an engine, and a gas turbine engine.
  • the gearbox includes a gear assembly coupled with the propulsor.
  • the engine includes an air inlet, an exhaust outlet, and an engine output shaft.
  • the engine output shaft is coupled with the gear assembly and configured to drive rotation of the propulsor through the gear assembly.
  • the gas turbine engine includes a first rotational assembly, a second rotational assembly, a compressor section, a turbine section, and a combustor. The first rotational assembly and the second rotational assembly are configured for rotation about a rotational axis.
  • the first rotational assembly includes a bladed compressor rotor for the compressor section, a bladed first turbine rotor for the turbine section, and a first shaft interconnecting the bladed compressor rotor and the bladed first turbine rotor.
  • the second rotational assembly includes a bladed second turbine rotor and a second shaft. The first shaft and the second shaft are operably coupled with the gear assembly to further drive rotation of the propulsor through the gear assembly.
  • the compressor section is connected to the air inlet and configured to direct a compressed air to the air inlet.
  • the combustor is connected to the exhaust outlet and configured to receive an exhaust gas from the exhaust outlet.
  • the combustor is configured to direct a combustion gas through the turbine section to drive rotation of the bladed first turbine rotor.
  • the bladed second turbine rotor may be downstream of the first bladed turbine rotor with respect to the combustion gas.
  • the second rotational assembly may be axially separated from the first rotational assembly.
  • the engine assembly may further include an electric control assembly including a first electric motor and an energy storage device.
  • the first electric motor may be coupled with the gear assembly.
  • the first electric motor may be electrically connected to the energy storage device.
  • the first electric motor may be configured to selectively apply a rotational force to the gear assembly to further drive rotation of the propulsor through the gear assembly.
  • the combustor may be disposed between the turbine section and the compressor section.
  • the engine assembly may further include a heat exchanger connected in fluid communication between the compressor section and the air inlet.
  • the heat exchanger may be configured to cool the compressed air directed from the compressor section to the air inlet.
  • an engine assembly for an aircraft propulsion system includes a propulsor, a gearbox, an engine, and a gas turbine engine.
  • the gearbox includes a gear assembly coupled with the propulsor.
  • the engine includes an air inlet, an exhaust outlet, and an engine output shaft.
  • the engine output shaft is coupled with the gear assembly and configured to drive rotation of the propulsor through the gear assembly.
  • the gas turbine engine includes a first rotational assembly, a second rotational assembly, a compressor section, a turbine section, and a combustor. The first rotational assembly and the second rotational assembly are configured for rotation about a rotational axis.
  • the first rotational assembly includes a bladed compressor rotor for the compressor section, a bladed first turbine rotor for the turbine section, and a first shaft interconnecting the bladed compressor rotor and the bladed first turbine rotor.
  • the second rotational assembly includes a bladed second turbine rotor and a second shaft. The second shaft is coupled with the gear assembly to further drive rotation of the propulsor through the gear assembly.
  • the compressor section is connected to the air inlet and configured to direct a compressed air to the air inlet.
  • the combustor is connected to the exhaust outlet and configured to receive an exhaust gas from the exhaust outlet. The combustor is configured to direct a combustion gas through the turbine section to drive rotation of the bladed first turbine rotor.
  • the engine assembly may further include an electric control assembly including a first electric motor and an energy storage device.
  • the first electric motor may be coupled with the gear assembly.
  • the first electric motor may be electrically connected to the energy storage device.
  • the first electric motor may be configured to selectively apply a rotational force to the gear assembly to further drive rotation of the propulsor through the gear assembly.
  • the electric control assembly may further include a second electric motor.
  • the second electric motor may be coupled with the first rotational assembly.
  • the second electric motor may be electrically connected to the energy storage device.
  • the engine assembly may further include a controller including a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, may cause the processor to control the second electric motor to generate a first electrical power during a first operating condition of the engine assembly and apply the rotational force to the gear assembly during the first operating condition of the engine assembly while the second electric motor generates the first electrical power.
  • a controller including a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, may cause the processor to control the second electric motor to generate a first electrical power during a first operating condition of the engine assembly and apply the rotational force to the gear assembly during the first operating condition of the engine assembly while the second electric motor generates the first electrical power.
  • the engine 28 of FIG. 2 is configured as an internal combustion engine using intermittent combustion during operation.
  • the engine 28 is an engine configuration other than a gas turbine engine configuration.
  • the engine 28 may be a reciprocating engine such as, but not limited to, a piston engine or a rotary engine (e.g., a Wankel engine).
  • the engine 28 includes an air inlet 36, an exhaust outlet 38, and an engine output shaft 40.
  • the engine 28 receives compressor air from the gas turbine engine 30 at the air inlet 36 and directs exhaust gas to the gas turbine engine 30 from the exhaust outlet 38, as will be discussed in further detail.
  • the engine 28 drives rotation of the engine output shaft 40.
  • the engine output shaft 40 is coupled with the gearbox 32 to drive, at least in part, rotation of the propulsor 22.
  • the gas turbine engine 30 of FIG. 2 includes a compressor section 42, a turbine section 44, and a combustor 46 arranged along a rotational axis 48 (e.g., an axial centerline) of the gas turbine engine 30.
  • the compressor section 42 of FIG. 2 includes a compressor 42A.
  • the turbine section 44 of FIG. 2 (e.g., a recovery turbine section) includes a high-pressure turbine 44A and a low-pressure turbine 44B (e.g., a power turbine).
  • the combustor 46 may be disposed between the compressor section 42 and the turbine section 44 along the rotational axis 48.
  • the compressor section 42 includes an air inlet 50 and an air outlet 52.
  • the air inlet 50 is connected in fluid communication with an air intake of the propulsion system 20 to receive a flow of ambient air from outside the propulsion system 20.
  • the air outlet 52 is connected in fluid communication with the air inlet 36 to direct compressed air from the compressor 42A to the air inlet 36.
  • the engine assembly 24 may include a heat exchanger 54 (e.g., an intercooler) connected in fluid communication between the air outlet 52 and the air inlet 36 to cool the compressed air directed to the air inlet 36 from the air outlet 52.
  • the heat exchanger 54 may be configured to receive ambient air (e.g., from a propulsion system 20 air scoop or another air source) for cooling the compressed air.
  • the present invention is not limited to any particular cooling medium for the heat exchanger 54.
  • Components of the compressor section 42 and the turbine section 44 form a first rotational assembly 56 (e.g., a high-pressure spool) and a second rotational assembly 58 (e.g., a low-pressure or power spool) mounted for rotation about the rotational axis 48.
  • first rotational assembly 56 e.g., a high-pressure spool
  • second rotational assembly 58 e.g., a low-pressure or power spool mounted for rotation about the rotational axis 48.
  • the gas turbine engine 30 is described herein as including a two-spool configuration, the present invention is also applicable to single-spool gas turbine engine configurations.
  • the first rotational assembly 56 includes a first shaft 60, a bladed compressor rotor 62 for the compressor 42A, and a bladed first turbine rotor 64 for the high-pressure turbine 44A.
  • the first shaft 60 interconnects the bladed compressor rotor 62 and the bladed turbine rotor 64.
  • the first shaft 60 is coupled with the electric control assembly 34, as will be discussed in further detail.
  • the second rotational assembly 58 includes a second shaft 66 and a bladed second turbine rotor 68 for the low-pressure turbine 44B.
  • the second shaft 66 interconnects the bladed second turbine rotor 68 with the gearbox 32.
  • the second shaft 66 is coupled with the gearbox 32 to drive, at least in part, rotation of the propulsor 22.
  • the first shaft 60 and the second shaft 66 of FIG. 2 are axially separated and rotate about the rotational axis 48.
  • the present invention is not limited to the particular rotational assembly configuration of FIG. 2 , and the first shaft 60 and the second shaft 66 may alternatively be concentric relative to the rotational axis 48.
  • the first shaft 60 and the second shaft 66 may also alternatively be configured for rotation about discrete rotational axes.
  • the gearbox 32 of FIG. 2 includes a gear assembly 70.
  • the gear assembly 70 is coupled with the engine drive shaft 40, the second shaft 66, and the electric control assembly 34 (e.g., an electric motor of the electric control assembly 34). Rotational inputs of the engine drive shaft 40, the second shaft 66, and the electric control assembly 34 are combined by the gear assembly 70 to drive rotation of the propulsor 22.
  • the gear assembly 70 may be additionally coupled with an input shaft 74 of the propulsor 22.
  • the gear assembly 70 may include an epicyclic gear assembly (e.g., a planetary gear assembly) coupling the engine drive shaft 40, the second shaft 66, and the electric control assembly 34 with the propulsor 22.
  • the present invention is not limited to any particular gear configuration for the gear assembly 70.
  • the electric control assembly 34 of FIG. 2 includes a first electric motor 76, a second electric motor 78, and an energy storage device 80.
  • the electric control assembly 34 may additionally include one or more inverters 82.
  • the first electric motor 76 includes a motor drive shaft 84 coupled with the gear assembly 70.
  • the first electric motor 76 is configured to apply a rotational force to the gear assembly 70 to drive, at least in part, rotation of the propulsor 22.
  • the first electric motor 76 may additionally be configured as a starter motor for the engine 28, for example, by applying a rotational force to the gear assembly 70 to drive rotation of the engine drive shaft 40.
  • the first electric motor 76 may be configured as an AC motor (e.g., a synchronous AC motor, an induction AC motor, etc.) or a DC motor.
  • the first electric motor 76 may additionally have a motor-generator (MG) configuration such that the first electric motor 76 may also generate electrical power for use by the second electric motor 78 and/or storage by the energy storage device 80.
  • the first electric motor 76 is electrically connected to the energy storage device 80.
  • the first electric motor 76 of FIG. 2 is electrically connected to the energy storage device 80, in part, by the inverter 82.
  • the inverter 82 is configured to control an electrical power supply (e.g., voltage, current, and frequency) to the first electric motor 76 to control operation (e.g., rotation speed, output torque, etc.) of the first electric motor 76.
  • an electrical power supply e.g., voltage, current, and frequency
  • the inverter 82 may be configured to convert direct current (DC) electrical power from the energy storage device 80 to alternating current (AC) electrical power for use by the first electric motor 76.
  • the inverter 82 may be further configured to convert AC electrical power generated by the first electric motor 76 to DC electrical power for storage by the energy storage device 80.
  • the second electric motor 78 is coupled with the first rotational assembly 56 (e.g., the first shaft 60).
  • the second electric motor 78 may be directly coupled with the first shaft 60.
  • the second electric motor 78 may be coupled with the first shaft 60 by a gearbox (e.g., a gear assembly), a clutch, or the like.
  • the second electric motor 78 is configured to apply a rotational force to the first shaft 60 to selectively drive, at least in part, rotation of the first rotational assembly 56.
  • the second electric motor 78 may be configured as an AC motor (e.g., a synchronous AC motor, an induction AC motor, etc.) or a DC motor.
  • the second electric motor 78 may additionally have a motor-generator (MG) configuration such that the second electric motor 78 may also generate electrical power for use by the first electric motor 76 and/or storage by the energy storage device 80.
  • the second electric motor 78 is electrically connected to the energy storage device 80.
  • the second electric motor 78 of FIG. 2 is electrically connected to the energy storage device 80, in part, by the inverter 82.
  • the inverter 82 is configured to control an electrical power supply (e.g., voltage, current, and frequency) to the second electric motor 78 to control operation (e.g., rotation speed, output torque, etc.) of the second electric motor 78.
  • an electrical power supply e.g., voltage, current, and frequency
  • the inverter 82 may be configured to convert direct current (DC) electrical power from the energy storage device 80 to alternating current (AC) electrical power for use by the second electric motor 78.
  • the inverter 82 may be further configured to convert AC electrical power generated by the second electric motor 78 to DC electrical power for storage by the energy storage device 80.
  • the first electric motor 76 and the second electric motor 78 may be electrically connected to a shared inverter 82, as shown in FIG. 2 , which shared inverter 82 may control the operation of the first electric motor 76 and the second electric motor 78.
  • the first electric motor 76 and the second electric motor 78 may be controlled by discrete inverters 82.
  • Operation of the engine 28 includes a first combustion stage (e.g., an intermittent internal combustion stage) for the engine assembly 24 in which a fuel (e.g., diesel fuel, kerosene fuel, etc.) is mixed and burned with the compressed air from the air inlet 36 and resultant combustion gas is exhausted from the engine 28 at the exhaust outlet 38 as exhaust gas 86.
  • the exhaust gas 86 is directed from the exhaust outlet 38 to the combustor 46.
  • Operation of the engine 28 drives rotation of the engine drive shaft 40 which, in turn, drives rotation of the propulsor 22 (e.g., through the gear assembly 70.
  • the first combustion stage may include a lean fuel-to-air ratio.
  • the compressed air directed to the engine 28 from the compressor 42A, and subsequently mixed with the fuel may be in excess of a stoichiometric ratio for the fuel-air mixture.
  • the exhaust gas 86 from the engine 28 may include a substantial quantity of oxygen which was unused in the first combustion stage. Because a significant amount of engine assembly 24 power may be used in the operation of the compressor 42A to compress the air supplied to the engine 28, this unused oxygen in the exhaust gas 86 could, in at least some conventional aircraft propulsion systems, represent wastage of engine assembly 24 power and, therefore, reduced engine assembly 24 efficiency.
  • the exhaust gas 86 is directed to the combustor 46 to facilitate a second combustion stage for the engine assembly 24 in which a fuel is mixed and burned within the exhaust gas 86 within a combustion chamber of the combustor 46.
  • the fuel directed to the combustor 46 for the second combustion stage may be the same as or different than the fuel used for the first combustion stage.
  • This second combustion stage within the combustor 46 is accomplished with the unused oxygen in the exhaust gas 86 produced in the first combustion stage in the engine 28.
  • additional engine power for the engine assembly 24 can be extracted by the gas turbine engine 30 using the unused oxygen in the exhaust gas 86.
  • the bladed second turbine rotor 68 drives rotation of the propulsor 22 with the second shaft 66 (e.g., through the gear assembly 70) in combination with the engine 28 (e.g., the engine drive shaft 40) and/or the first electric motor 76.
  • the combustion gas 88 is exhausted from the propulsion system 20.
  • rotation of the first shaft 60 may drive the second electric motor 78 to generate electrical power and the electrical power generated by the second electric motor 78 may be used by the first electric motor 76 to apply rotational force to the gear assembly 70 to drive, at least in part, rotation of the propulsor 22.
  • the electrical power generated by the second electric motor 78 may be stored by the energy storage device 80.
  • the controller 26 may include, or may be in communication with, an input device that enables a user to enter data and/or instructions, and may include, or be in communication with, an output device configured, for example to display information (e.g., a visual display or a printer), or to transfer data, etc. Communications between the controller 26 and other electrical and/or electronic components (e.g., controllers, sensors, etc.) may be via a hardwire connection or via a wireless connection.
  • portions of the controller 26 may assume various forms (e.g., digital signal processor, analog device, etc.) capable of performing the functions described herein.
  • the EEC may modulate the fuel flow using a closed-loop process in which an output power or other operating parameter of the engine assembly 24 is measured and fuel flow to the engine 28 and/or the gas turbine engine 30 is increased or decreased as a function of the measured output power or operational parameter.
  • the EEC may be part of a full authority digital engine control (FADEC) system for the propulsion system 20.
  • FADEC full authority digital engine control
  • the target rotation speed may be identified, for example, using predetermined target rotation speed values (e.g., look-up tables or other databases) stored in memory 92 for the air flow parameters measured using the sensors 104.
  • an output power of the first rotational assembly 56 may be greater than work required by the compressor 42A to compress ambient air for the engine 28 (e.g., the first rotational assembly 56 rotation speed may be greater than the target rotation speed).
  • the controller 26 may control electric control assembly 34 to selectively apply at least some rotational force from the first rotational assembly 56 to the propulsor 22.
  • output power of the first rotational assembly 56 may be variably extracted by the electric control assembly 34 and applied for driving rotation of the propulsor 22 (e.g., in combination with the engine 28 and the second rotational assembly 58), thereby reducing engine 28 loading (e.g., fuel use) while maintaining the desired bladed compressor rotor 62 and propulsor 22 rotation speeds.
  • engine 28 loading e.g., fuel use
  • any one of these structures may describe the operations as a sequential process, many of the operations can be performed in parallel or concurrently.
  • the order of the operations may be rearranged.
  • a process may correspond to a method, a function, a procedure, a subroutine, a subprogram, etc.
  • the terms “comprise”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Engine Equipment That Uses Special Cycles (AREA)
  • Supercharger (AREA)
EP24200082.6A 2023-09-12 2024-09-12 Ensemble moteur pour système de propulsion d'aéronef Pending EP4524384A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US18/367,286 US12162613B1 (en) 2023-09-12 2023-09-12 Engine assembly for an aircraft propulsion system

Publications (1)

Publication Number Publication Date
EP4524384A1 true EP4524384A1 (fr) 2025-03-19

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EP24200082.6A Pending EP4524384A1 (fr) 2023-09-12 2024-09-12 Ensemble moteur pour système de propulsion d'aéronef

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EP (1) EP4524384A1 (fr)
CA (1) CA3254257A1 (fr)

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KR20250081978A (ko) * 2023-11-28 2025-06-09 한국항공우주연구원 다중 동력원 전기추진시스템을 구비한 비행체
US12583596B2 (en) * 2024-02-07 2026-03-24 Pratt & Whitney Canada Corp. Aircraft propulsion system
US12576979B2 (en) * 2024-02-29 2026-03-17 Pratt & Whitney Canada Corp. Aircraft propulsion system with engine assembly, intercooler, and turbocompressor assembly

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US20100219779A1 (en) * 2009-03-02 2010-09-02 Rolls-Royce Plc Variable drive gas turbine engine
US8141360B1 (en) * 2005-10-18 2012-03-27 Florida Turbine Technologies, Inc. Hybrid gas turbine and internal combustion engine
US20200298988A1 (en) * 2019-03-18 2020-09-24 Pratt & Whittney Canada Corp. Architectures for hybrid-electric propulsion

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US7980052B1 (en) 2010-05-20 2011-07-19 Florida Turbine Technologies, Inc. Industrial gas turbine engine
US9869240B2 (en) * 2015-02-20 2018-01-16 Pratt & Whitney Canada Corp. Compound engine assembly with cantilevered compressor and turbine
US11623757B2 (en) * 2018-06-08 2023-04-11 Embraer S.A. Hybrid electric taxi system (HETS) or full electric taxi system (FETS)
US11866181B2 (en) * 2021-09-15 2024-01-09 Pratt & Whitney Canada Corp. Aircraft power plant
US20240151179A1 (en) * 2022-11-04 2024-05-09 Raytheon Technologies Corporation Compounded turbo power unit with boost combustor

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Publication number Priority date Publication date Assignee Title
US8141360B1 (en) * 2005-10-18 2012-03-27 Florida Turbine Technologies, Inc. Hybrid gas turbine and internal combustion engine
US20100219779A1 (en) * 2009-03-02 2010-09-02 Rolls-Royce Plc Variable drive gas turbine engine
US20200298988A1 (en) * 2019-03-18 2020-09-24 Pratt & Whittney Canada Corp. Architectures for hybrid-electric propulsion

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CA3254257A1 (fr) 2025-05-30
US12162613B1 (en) 2024-12-10

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